JP4840664B2 - Method for compensating blade tip clearance degradation in active clearance control - Google Patents

Description

  The present invention relates to a method of compensating for turbine blade tip clearance degradation in a gas turbine engine.

  Engine performance parameters such as thrust, fuel consumption rate (SFC) and exhaust gas temperature (EGT) margin are highly dependent on the gap between the turbine blade tip and the stationary seal or shroud surrounding the blade tip. Over the life of the engine, these gaps tend to increase as a result of blade friction, oxidation, and erosion, reducing engine performance. Compensating for this drop (degradation) is highly desirable.

  It is well known that the main factor in reducing engine efficiency and increasing fuel consumption in aircraft gas turbine engines is a gradual increase in the gap between the turbine blade tip and the surrounding stationary seal or shroud. The tip clearance degradation increases the amount of turbine working fluid leakage through the individual turbine rotor stages and compressor stages of the gas turbine engine. Such leakage reduces the total engine efficiency and thus increases the total fuel consumption rate.

  This increase in tip clearance is directly related to the engine's cumulative usage since the engine was first used, or after engine maintenance has been performed and the tip clearance has been restored to or near official specifications. Yes. Engine maintenance generally involves the replacement or modification of seals or blade tips and requires time consuming and expensive operations. Immediately after installation, the blade tip clearance is minimal and the sealing effect is maximal. Blade tip clearance and sealing effects degrade as the engine is operated over an ever increasing cycle.

  One way to improve both the wear life and effectiveness of the blade tip seal is “active gap control”. Active clearance control is from engine fans and / or compressors that inject onto a high pressure turbine casing under specific operating conditions, i.e. steady state high altitude cruising conditions, causing the casing to contract against the high pressure turbine blade tips. Adjust the flow rate of cold air. Cooling air can flow or be jetted onto the stationary structure used to support the shroud or seal around the blade tip. The regulated cooling air flow rate is used under certain engine operating conditions and reduces or eliminates interference or friction between the seal and blade tips that may occur during transient conditions such as takeoff, deceleration, etc. However, it is designed to allow the engine to be operated with minimal seal clearance during the majority of its operating cycle.

Engines with an active clearance controller still produce blade tip clearance degradation related to cumulative engine usage. Attempts are known to measure the working blade tip clearance and attempt to correct the clearance by either mechanical or thermal means. It is also known to periodically recover the blade tip clearance by gradually increasing the flow rate of external cooling air in response to only cumulative engine usage parameters. Such a method is disclosed in US Pat. No. 4,856,272. This method sets incremental changes in the engine's active clearance control system at predetermined engine usage intervals between scheduled overhauls. Incremental and usage intervals are predetermined by past experience with a plurality of similarly configured engines, allowing adjustments to be made for a particular engine based solely on the accumulated time or cycle of engine operation.
U.S. Pat. No. 4,856,272 US Patent Application Publication No. 2005 / 0017876A1 US Patent Application Publication No. 2005 / 0149274A1 U.S. Pat. No. 4,304,093 U.S. Pat. No. 4,928,240 US Pat. No. 5,012,420 US Pat. No. 5,081,830 US Pat. No. 6,463,380

  It is highly desirable to be able to maintain or restore the optimum blade tip clearance in an aircraft gas turbine engine between seals and / or blade tip replacements or modifications. It is also highly desirable to accurately and automatically compensate for engine performance degradation due to increased blade tip clearance due to wear.

  A method for compensating for blade tip clearance degradation between a rotating blade tip and a surrounding shroud in an aircraft gas turbine engine includes at least one of one or more engine operating parameters averaged over a fixed number of operating engine flight cycles, or Determining one or more variables based on further moving average values, calculating a blade tip clearance degradation value based on the variables, and blade tip clearance based on the one or more variables. Adjusting the flow rate of the thermal control air so as to cancel out the deterioration value. The engine operating parameters can be selected from the group including number of operating engine cycles, takeoff and cruise exhaust gas temperature, cruise turbine efficiency, takeoff and cruise maximum turbine speed, and cruise fuel flow.

  The method further includes determining a blade tip closure amount to offset the blade tip clearance degradation value, and the thermal control air flow rate when the closure amount matches an incremental criterion for adjusting the thermal control air flow rate. Adjusting. The variable can include a difference between a moving average value of the engine operating parameter and a corresponding baseline. The method further includes determining a percentage of the variable due to the radial blade tip clearance between the blade tip and the shroud by a percentage function that relates the percentage to the variable, and then determining the component degradation value as the component degradation value. Determining by means of a component degradation function relating to the percentage.

  The above aspects and other features of the invention will become apparent from the following description taken in conjunction with the accompanying drawings.

  Shown schematically in cross-section in FIG. 1 is an exemplary embodiment of an aircraft gas turbine engine 10 that includes an active clearance control system 12. Based on one or more moving average values MAVG of one or more functions F of the engine performance parameter P, the active clearance control system 12 determines the blade tip clearance degradation value DT as a function of engine cycle in FIG. A method of compensating for Engine 10 is in a series flow relationship with fan section 13 with fan 14, booster or low pressure compressor (LPC) 16, high pressure compressor (HPC) 18, combustion section 20, high pressure turbine (HPT) 22, and low pressure turbine. (LPT) 24. A high pressure shaft 26 disposed around the engine centerline 8 drivably connects the HPT 22 to the HPC 18, and a low pressure shaft 28 drivably connects the LPT 24 to the LPC 16 and the fan 14. The HPT 22 includes an HPT rotor 30 that has a turbine blade 34 attached to its periphery.

  Shown in FIG. 2 is a stator assembly 64 mounted to the radially outer casing 66 of the HPT 22 by front and rear case hooks 68 and 70. The stator assembly 64 includes an annular segmented stator shroud 72 attached to an annular segmented shroud support 80 by front and rear shroud hooks 74 and 76. The shroud 72 surrounds the turbine blades 34 of the rotor 30 and serves to reduce leakage of flow around the radially outer blade tips 82 of the blades 34. The active clearance control system 12 is used to minimize the radial blade tip clearance CL between the blade tip 82 and the shroud 72 as shown in FIG.

  It is well known in the industry that a small turbine blade tip clearance CL lowers the operating fuel consumption rate SFC, thus resulting in significant fuel savings. Front and rear thermal control rings 84 and 86 are provided to more effectively control the blade tip clearance CL with a minimum amount of time delay and thermal control (cooling or heating depending on operating conditions) air flow rate. Front and rear thermal control rings 84 and 86 are associated with outer casing 66 and are either integral with the respective casing (as shown in FIG. 2) or bolted or otherwise secured to the casing. Alternatively, it can be mechanically separated from the casing but in sealing engagement with the casing. The thermal control ring constitutes a thermal control mass to move the stator shroud 72 more radially inward (and outward if so designed) to reduce the blade tip clearance CL. adjust. The injection tube 60 impinges the stator shroud 72 radially inward by impinging the thermal control air 36 (cooling air) on the front and rear thermal control rings 84 and 86 and, if so designed, the outer casing 66. To close or minimize the blade tip clearance CL.

  With reference to FIGS. 1 and 2, the compressed fan air supply 32 can be used as a source for the thermal control air 36, and the thermal control air 36 is entirely connected via an axial air supply pipe 42. Is supplied to a turbine blade tip clearance control device indicated by reference numeral 40. The air supply inlet 19 to the axial air supply pipe 42 is installed downstream of the outlet guide vane 17 disposed in the fan bypass duct 15 downstream of the fan 14. The air valve 44 disposed in the air supply pipe 42 controls the total amount of the heat control air 36 that flows in the air supply pipe 42. The thermal control air 36 is cooling air in this exemplary embodiment of the active gap control system 12 shown herein. Cooling air is controllably flowed from the fan bypass duct 15 surrounding the booster or low pressure compressor (LPC) 16 through the axial air supply pipe 42 to the distribution manifold 50 of the turbine blade clearance controller 40.

  The amount of thermal control air 36 that impinges to control the air valve 44 and the turbine blade tip clearance CL (shown in FIG. 2) is controlled by the controller 48. The controller 48 is a digital electronic engine control system, often referred to as a fully automatic digital electronic control unit (FADEC), which controls the amount and temperature of the thermal control air 36 impinging on the front and rear thermal control rings 84 and 86 as needed. Therefore, the turbine blade tip clearance CL is controlled. Manifold 50 includes an annular header tube 54 that distributes cooling air to a plurality of plenums 56, which then distributes the cooling air to a plurality of injection tubes 60, as shown in FIG.

  An algorithm, referred to below as an ACC flow model 92, a mathematically computed active clearance control flow model, is used to control the turbine blade tip clearance CL, stored in the controller 48 and executed in the controller 48. The ACC flow model 92 is based on the engine operating parameters of the engine and the physical characteristics of various components. The controller 48 sends a valve position signal to the air valve 44 based on the calculated ACC flow model 92 to control the total amount of thermal control air 36. The air valve 44 is gradually opened according to the valve position signal. The ACC flow model 92 is based at least in part on the calculated amount of blade tip degradation DT. As shown in FIG. 9, the radial blade tip clearance CL increases the amount of blade tip degradation DT as the engine is used and the amount of cycle increases. In this exemplary embodiment shown herein, the ACC flow model 92 includes an additional degradation term to account for the amount of blade tip clearance degradation DT. A gap model program CLM for determining the amount of blade tip degradation DT is executed in FADEC as part of the ACC flow model 92. As shown in the flowchart of FIG. 3, the gap model program CLM is executed as a background of FADEC after the engine is started.

  The calculated turbine blade tip clearance degradation value DT is offset by the turbine blade tip closing value TCL. The calculated turbine blade tip clearance degradation value DT is one or more engine operating parameters P, each as shown in FIG. 4 by a graph of DT versus function F of MAVG (I) expressed as F (MAVG (I)). Based on one or more moving average values MAVG (I) of (I). I represents an engine operation parameter index with respect to the number of different engine operation parameters P being used. For example, when five different engine operation parameters are used, I = 1 to 5.

  The turbine blade tip closure value TCL and the calculated turbine blade tip clearance degradation value DT are each one or more of one or more moving average values MAVG (I) of one or more engine operating parameters P (I). Based on the above function F (I). The moving average value is averaged over a period that is a constant operating engine flight cycle number NC. This exemplary method presented herein uses 50 engine cycle periods as shown in FIG. The flow rate of the thermal control air 36 is adjusted so that the turbine blade tip closing value TCL cancels the blade tip deterioration value DT. The tip clearance deterioration value DT is determined by a clearance deterioration compensation model 94 stored in the controller 48 and executed in the controller 48 as shown in FIG.

  Engine operating parameters may include number of operating engine cycles, takeoff and cruise exhaust gas temperature EGT, cruise turbine efficiency, takeoff and cruise maximum turbine speed N, cruise fuel flow rate, and others. This method may use the difference D between the moving average value MAVG of one or more engine operating parameters P and the respective baseline B of the one or more engine operating parameters. Baseline B may be selected from existing values based on experimental, semi-empirical or analytical methods or based on values measured or determined during engine operation as in the methods described herein. Herein, the moving average is shown as averaged over the first 50 engine cycle periods. The baseline herein is also a moving average value averaged over the first 50 engine cycle periods as shown in FIG.

  Thus, the calculated tip clearance degradation value DT can be based on what is referred to herein as one or more variables V, which are one of one or more engine operating parameters P. Or a difference D between the moving average value MAVG and / or the moving average value MAVG of one or more engine operating parameters P and the respective baseline B of the one or more engine operating parameters P. Or a combination of the two. V (I) can be an engine operating parameter P such as the total number of engine cycles. Therefore, if I = 1 to N (where N is the number of engine operating parameters P used in the method for determining the calculated tip clearance degradation value DT as shown in FIG. 5), V (I ) = D (I) or MAV (I) or P (I). If one V (I) is P (I) such as an engine cycle, each of one or more moving average values MAVG or moving average values MAVG and one or more engine operating parameters P At least one other V (I) based on the difference D (I) from the baseline B is used.

  This exemplary method for calculating the blade tip clearance degradation value DT shown herein uses a computed composite turbine blade tip clearance degradation DTC model stored in and operating within the FADEC. The clearance model program CLM for determining the amount of blade tip degradation DT, exemplified by the calculated composite turbine blade tip clearance degradation value DTC, is executed in FADEC as part of the ACC flow model 92. The composite turbine blade tip clearance degradation value DTC is composed of a component degradation value DCL (I), where I represents different engine operating parameters shown as DCL (1) through DCL (7) in FIG. The grand total or composite turbine blade tip clearance degradation value DTC is calculated by first determining all component degradation values DCL (I) and then summing all of the component degradation values DCL (I). There is a different component degradation value DCL (I) for each of the one or more engine operating parameters P (I).

  Shown in FIG. 7 are some specific variables V of the engine operating parameter P and one or more moving average values of the cycle. One of the variables V is due to the engine cycle, and the rest of the variable V is a moving average value of one or more engine operating parameters P and a respective baseline of the one or more engine operating parameters. This is due to the difference D from B. The difference D shown in FIG. 7 represents the moving average values for takeoff and cruise exhaust gas temperatures DTEGT and DCEGT, cruise turbine efficiency DTEFF, takeoff and cruise maximum turbine speeds DTN and DCN, and cruise fuel flow rate DWF and their respective baselines. B.

  The component deterioration value DCL (I) is based on one or more moving average values of at least one or more engine operating parameters P (I), respectively. This exemplary method presented herein is based on the component degradation value DCL (I) being based on a number of variables V (I), which in turn are the engine operating parameters P (I ) Based on one or more moving average values. One of the variables V (I) is a moving average value of the engine cycle, and the rest of the variable V (I) is one or more moving average values of the engine operating parameter P (I) and the one or The difference D (I) between the respective baselines of further engine operating parameters. The difference D (I) is the take-off and cruise exhaust gas temperatures DTEGT and DCEGT, cruise turbine efficiency DTEFF, take-off and cruise maximum turbine speeds DTN and DCN, and cruise fuel flow rate DWF baseline B (I) and their respective moving average values. It is the difference between MAVG (I).

  This exemplary method presented herein utilizes a two-step process performed by the active gap control flow model ACC flow model 92 stored in the FADEC to determine the component degradation value DCL (I). The first stage, shown as stage 1 in FIG. 6, determines the percentage% V (I) of the variable V (I) due to the radial blade tip clearance CL between the blade tip 82 and the shroud 72. In determining the percentage% V (I), the percentage function% F (I) stored as a function of the variable V (I) in the controller 48 or FADEC is used. Many methods are known for storing functions in FADEC, and the method shown here is a table. A table lookup method or scheme is used to determine the value of the percentage function% F (I) based on the respective variable V (I). Table lookup methods typically store values as a function of variables and use various interpolation schemes to produce an output value such as% V (I) for any input value such as V (I). decide.

  The second stage shown as stage 2 in FIG. 6 determines the component degradation value DCL (I), which is the component degradation function FDCL of the percentage% V (I) determined in stage 1. Stored as (I). The component deterioration value DCL (I) is stored in the controller 48 or FADEC. Many methods are known for storing functions in FADEC, the method shown here is a table, and the table lookup method or scheme is the percentage function% F (I for each variable V (I). ) Is used to determine the component deterioration value DCL (I). A table lookup method typically stores values as a function of variables and uses various interpolation schemes to produce an output value such as DCL (I) for any input value such as% V (I). decide. The component degradation value DCL (I) is then summed to calculate a composite turbine blade tip clearance degradation value DCL.

  Prior to summing the component degradation values DCL (I) to calculate the composite turbine blade tip clearance degradation value DCL, a confidence or weighting factor CF (I) can be applied to the component degradation value DCL (I). When reliability or weight coefficient CF (I) is used for K engine operating parameters P (I), composite turbine blade tip clearance degradation value DCL = (CF (1) * DCL (1) + CF (2) * DCL (2) + ... CF (K) * DCL (K)) / (CF (1) + CF (2) + ... CF (K)). Since the moving average value is averaged over a predetermined engine cycle number NC or a moving average period, at least after the predetermined engine cycle number NC, no change is made to the turbine blade tip baseline gap CLB. The turbine blade tip clearance baseline is the nominal cold clearance of the assembly, typically based on part drawing dimensions.

  The functions F (I) exemplified in F (1) to F (7) in FIG. 7 are in the form of algorithms or tables, and the component blade tip gap deterioration value DCL (I) is determined by these algorithms or tables, It is stored in the controller 48. The engine operating parameters shown in this specification are inputs for determining the high pressure turbine efficiency HPTEFF, fuel flow rate WF, takeoff and cruise exhaust gas temperature EGT, takeoff and cruise maximum value N2 (rotational speed of the high pressure turbine). The take-off engine operating parameter value is measured and recorded during take-off, while the cruise engine operating parameter value is stable cruise state, for example about 10 minutes after reaching the cruise altitude during the flight cycle as shown in FIG. Is recorded.

  In this exemplary embodiment of the method presented herein, the air valve 44 is incrementally increased according to the valve position signal so that the calculated turbine blade tip clearance degradation value DT is offset by incrementally increasing the turbine blade tip closure value TCL. The turbine blade tip closure value TCL is shown herein as gradually increasing (incrementing) by 0.005 inches as shown in FIG. The air valve 44 is opened incrementally, thus gradually increasing the total amount of flowing thermal control air 36 and increasing the amount of turbine blade tip closure TCL. An exemplary incremental amount of increase in turbine blade tip closure value TCL is shown herein as 0.005 inches.

  In this exemplary embodiment of the method presented herein, the air valve 44 is opened only when the calculated turbine blade tip closure value DT is within the predetermined and calculated upper and lower limits UB and LB, respectively. These upper and lower limits UB and LB can be based on the number of operating engine cycles, as shown in FIG. 9, respectively, in this exemplary embodiment shown herein.

  While this specification has described what is considered to be preferred and exemplary embodiments of the invention, other modifications of the invention will become apparent to those skilled in the art from the teachings herein. Therefore, all such modifications are desired to be protected by the following claims as falling within the spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent is the invention as defined and identified by the following claims.

1 is a schematic cross-sectional view of an aircraft gas turbine engine active clearance control system employing a blade tip clearance degradation compensation method based on a moving average value of engine performance parameters. FIG. 2 is a cross-sectional view of a blade gap control device used in the active gap control system shown in FIG. 1. FIG. 2 is a flowchart showing a method for determining a blade tip deterioration value DT for a blade tip gap deterioration compensation method used in the active gap control system shown in FIG. 1. FIG. 2 is a graph schematically showing a relationship between a blade tip gap closing component for compensating for the deterioration shown in FIG. 1 and a moving average value of engine operating parameters. FIG. 3 is a graph illustrating another exemplary relationship between a blade tip clearance closure component and a moving average value of engine operating parameters to compensate for the degradation shown in FIG. 1. FIG. 5 is a two-stage graph illustrating an exemplary relationship between a blade tip clearance closure component and a moving average value of engine operating parameters to compensate for the degradation shown in FIG. FIG. 3 is a set of graphs illustrating another exemplary relationship between an exemplary component of blade tip clearance closure to compensate for the degradation illustrated in FIG. 1 and a moving average value of exemplary engine operating parameters. FIG. 3 is an enlarged sectional view of a blade tip gap and deterioration shown in FIG. 2. The graph which shows the predetermined upper limit and lower limit of the blade tip deterioration term used in the active gap control flow model shown in FIG.

Explanation of symbols

8 Engine centerline 10 Aircraft gas turbine engine 12 Active clearance control system 13 Fan section 14 Fan 15 Fan bypass duct 16 Booster or low pressure compressor (LPC)
17 Exit guide vane 18 High pressure compressor (HPC)
19 Air supply inlet 20 Combustion section 22 High pressure turbine (HPT)
24 Low pressure turbine (LPT)
26 High-pressure shaft 28 Low-pressure shaft 30 HPT rotor 32 Fan air supply device 34 Turbine blade 36 Thermally controlled air 40 Turbine blade tip clearance control device 42 Air supply pipe 44 Air valve 48 Controller 50 Distribution manifold 54 Header pipe 56 Plenum 60 Injection pipe 64 Stator Assembly 66 Outer casing 68 Front case hook 70 Rear case hook 72 Shroud 74 Front shroud hook 76 Rear shroud hook 80 Shroud support 82 Blade tip 84 Front heat control ring 86 Rear heat control ring 92 ACC flow model 94 Gap deterioration compensation model CL Blade tip clearance DT Blade tip clearance degradation value B Baseline D Difference F Function P Parameter V Variable NC Flight cycle / Engine cycle UB Top LB lower CLB tip baseline clearance value CLM gap model program EGT exhaust gas temperature SFC fuel consumption TCL turbine blade tip closure value

Claims (10)

A method of compensating for blade tip clearance degradation (DT) between a rotating blade tip (82) and a surrounding shroud (72) in an aircraft gas turbine engine (10) comprising:
Determining one or more variables (V) based on at least one or more moving average values of one or more engine operating parameters, each averaged over a fixed number of operating engine flight cycles; ,
Calculating a blade tip clearance degradation value (DT) based on the variable (V);
Adjusting the flow rate of thermal control air (36) to offset the blade tip clearance degradation value based on the one or more variables;
Including methods.   One or more of the engine operating parameters selected from the group comprising operating engine cycle number, takeoff and cruise exhaust gas temperature (EGT), cruise turbine efficiency, takeoff and cruise maximum turbine speed and cruise fuel flow, The method of claim 1. Determining a blade tip closure amount to offset the blade tip gap degradation value;
Adjusting the flow rate of the thermal control air (36) when the closure amount matches an incremental criterion for adjusting the flow rate of the thermal control air (36);
The method of claim 2 further comprising:   The variable is a difference between one or more moving average values of the one or more engine operating parameters and one or more baselines of the one or more engine operating parameters, respectively. The method of claim 1 comprising:   The blade tip gap deterioration value (DT) is calculated from a corresponding variable of the variable (V (I)), and a composite turbine blade tip gap deterioration value (DTC) including a component deterioration value (DCL (I)) is calculated, and The method of claim 4, further comprising calculating by summing the components to determine a blade tip clearance degradation value (DT). The percentage (% V (I)) of the variable (V (I)) due to the radial blade tip clearance (CL) between the blade tip (82) and the shroud (72) is expressed as the percentage (% V (I )) By a percentage function (% F (I)) relating to the variable (V (I));
Next, the component deterioration value (DCL (I)) is determined by a component deterioration function (FDCL (I)) that relates the component deterioration value (DCL (I)) to the percentage (% V (I)). Stages,
The method of claim 5, further comprising: Determining a blade tip closure amount to offset the blade tip gap degradation value;
The closing amount coincides with an increment criterion for adjusting the flow rate of the thermal control air (36), and the blade tip clearance deterioration value is within a predetermined upper limit and lower limit range based on the number of operating engine cycles. Adjusting the flow rate of the thermal control air (36) when
The method of claim 6 further comprising:   The variable is a difference between a first portion of one or more moving average values of the one or more engine operating parameters and a corresponding baseline of the first portion, respectively; The method of claim 1, comprising a combination of one or more moving average values with a second portion.   The blade tip gap deterioration value (DT) is calculated from a corresponding variable of the variable (V (I)), and a composite turbine blade tip gap deterioration value (DTC) including a component deterioration value (DCL (I)) is calculated, and The method of claim 8, further comprising calculating by summing the components to determine a blade tip clearance degradation value (DT). The percentage (% V (I)) of the variable (V (I)) due to the radial blade tip clearance (CL) between the blade tip (82) and the shroud (72) is expressed as the percentage (% V (I )) By a percentage function (% F (I)) relating to the variable (V (I));
Next, the component deterioration value (DCL (I)) is determined by a component deterioration function (FDCL (I)) that relates the component deterioration value (DCL (I)) to the percentage (% V (I)). When,
The method of claim 9, further comprising:

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